There are known aircraft equipped with two engines mounted on the rear fuselage such as those shown in FIGS. 1a, 1b, 2a, 2b and 3.
FIGS. 1a and 1b show an aircraft with two turbofan engines 13 attached to the rear fuselage 11 by means of pylons 17 and an empennage comprising a vertical tail plane 21 and an upper horizontal tail plane 23 behind the propulsion system.
FIGS. 2a and 2b show an aircraft with two turboprop engines 13 attached to the rear fuselage 11 by means of pylons 17 and an empennage comprising a vertical tail plane 21 and an upper horizontal tail plane 23 behind the propulsion system.
FIG. 3 shows an aircraft with two turbofan engines 13 attached directly to the rear fuselage 11 and an empennage comprising a vertical tail plane 21 and an upper horizontal tail plane 23 behind the propulsion system.
In these aircraft, failure events such as a Blade Release (BR) event, i.e. an event where an external blade of one turboprop engine comes off and hits the fuselage, or an Uncontained Engine Rotor Failure (UERF) event, i.e. an event where a part of the internal rotors of the engine breaks, it is released and hits the fuselage, can generate large damages on the fuselage and also in the opposite engine. In the last case, the effects can be catastrophic.
Although engine manufacturers are making efforts to reduce the probability of said failure events, experience shows that UERF and BR events that can lead to catastrophic events continue to occur.
The certification requirements are very restrictive and are driving both fuselage and systems architectures in order to fulfill safety requirements.
As is well known, weight is a fundamental aspect in the aeronautic industry and therefore there is a trend to use structures of a composite material instead of a metallic material even for primary structures such as fuselages.
The composite materials that are most used in the aeronautical industry consist of fibers or fiber bundles embedded in a matrix of thermosetting or thermoplastic resin, in the form of a preimpregnated or “prepreg” material. Their main advantages refer to:                Their high specific strength with respect to metallic materials. It is the strength/weight equation.        Their excellent behavior under fatigue loads.        The possibilities of structural optimization thanks to the anisotropy of the material and the possibility of combining fibers with different orientations, allowing the design of the elements with different mechanical properties to be adjusted to the different needs in terms of applied loads.        The disadvantage of the usual composite materials made of carbon fibers compared to conventional light weight metallic materials like the aluminum, is their lower impact resistance and damage tolerance capabilities. No plasticity behavior as on metallic materials is present in composite materials and they are not able to absorb high strain energy amounts when deforming.        
There is therefore a need of fuselage structures capable to satisfy the safety requirements particularly when they are made up of composite materials.
Some proposals of impact resistant and damage tolerant fuselages are known in the prior art, which are capable of maintaining enough torsional strength when a part of the fuselage is removed as a consequence of one of said engine failure events for proceeding to the so called “get home mission with, only, the undamaged engine, such as those disclosed on WO 2009/068638 and US 2011/233335.
However none of the above-mentioned proposals can efficiently protect an engine (including systems such as electrical generation and fuel feed that are critical ones) from the risk of being damaged by a detached part from the opposite engine.
The present invention is addressed to the solution of this drawback.